A rotor blade spar is the foremost structural element of a helicopter rotor blade assembly inasmuch as its primary function is to transfer combined flapwise, edgewise, torsional and centrifugal loads to/from a central torque drive hub member. Typically, a leading edge sheath and trailing edge pocket assembly mount to and envelop the spar thereby yielding the desired airfoil contour. The spar typically extends the full length of the rotor blade and mounts at its inboard end to a cuff assembly or fitting which facilitates mounting to the hub member. Due to the extreme operational loading environment of the rotor blade, high strength, high density materials such as aluminum or titanium have, in the past, been the materials of choice for spar construction.
More recently, however, fiber reinforced resin matrix composite materials, e.g., graphite and fiberglass, have been employed due to their advantageous strength to weight ratio, corrosion resistance, and improved damage tolerance. Regarding the latter, the structural fibers of composite materials can be viewed as a plurality of redundant load paths wherein damage to one or more fibers can be mitigated by the load carrying capability of adjacent fibers.
Despite the inherent weight and strength advantages of advanced composites, the widespread use thereof has been impeded by the high cost of associated fabrication methods. Blending the desired structural characteristics with a low cost manufacturing process, i.e., one which reduces labor intensive process steps yet maintains laminate quality, has been an ongoing and continuous challenge for designers of composite structures.
Primary structural items to be considered by the designer include: the selection of fiber reinforcement, i.e., materials having the requisite mechanical properties, resin binder, fiber matrix orientation, fiber continuity, alleviation of stress concentrations due to ply drop-offs or joint configurations, and reduction of thermally induced stresses. To maximize the benefits of composites it is essential that the fiber orientation be optimally tailored to meet the strength and stiffness requirements for a particular application. That is, composites can be tailored to be anisotropic (capable of carrying load in a particular direction) rather than quasisotropic (equal strength in all directions); hence, orienting the fibers in the direction of the load will optimally result in the most weight efficient structure. Similarly, by varying the use of available matrix reinforcement materials (e.g., graphite, fiberglass, aramid fibers), the designer is able to control such parameters as vibratory and steady bending strength, stiffness, and toughness. In addition to the selection of materials and/or optimum fiber orientation, the continuity or discontinuity of fibers, and methods of joining discontinuous plies will significantly impact component strength. Generally, it is desirable to maintain fiber continuity and stagger joints so as to prevent stress concentrations and/or the build-up thereof in a particular region. Still other considerations relate to the thermal induced stresses which may result in microcracking. Microcracking is a phenomena wherein thermally induced stresses cause small cracks to develop in the binder material due to the thermal incompatibility of adjacent composite material. Generally, it is preferable to use the same material throughout the laminate or materials of similar thermal coefficient to reduce this effect.
These considerations are weighed and balanced against the cost and complexity of a particular fabrication technique. Typically, the manufacturing approach should: minimize cutting operations and material scrap, facilitate ease of handling, facilitate automation to a maximum practical extent, minimize the probability of operator/laminator error, be repeatable, maintain uniform fiber volume, provide uniform laminate quality (via uniform compaction), and yield a mass-balanced lay-up, i.e., properly distribute the weight of the composite article.
Conventional methods for manufacturing composite spars include filament winding and prepreg lay-up of composite material. The filament winding process involves the winding of high strength filaments disposed in a matrix of binder material about a mandrel assembly of a shape generally corresponding to the required shape of the finished article. The mandrel assembly typically comprises a rigid substructure to support the wound matrix and an impervious bladder or bag disposed over the rigid substructure. In the case of the manufacture of an elongated article, such as a rotor blade spar, the filaments are wound over the mandrel, and the mandrel and filaments are reciprocally displaced relative to one another along the longitudinal or winding axis of the mandrel to build a plurality of layers of filamentary material. Upon completion of the filament winding process, the mandrel/wound lay-up is placed in a matched metal mold and cured. During the curing process the bladder is pressurized to urge the fibers against the mold surfaces of the matched metal mold.
Advantages of filament winding include continuity of fibers (i.e., the process requires no cutting or ply overlaps), ease of handling, and repeatability. A primary disadvantage to filament winding, however, relates to difficulties associated with expanding/urging the fibers against the mold surfaces of a matched metal mold. Since the fibers are initially wound about the mandrel under tension, it is difficult to force the elongation and/or shifting of fibers during bag pressurization, to achieve proper laminate compaction. Should the fibers resist complete and uniform compaction, the composite article may become resin-rich or resin-starved in particular areas resulting in poor laminate quality. In the case of an elliptically shaped composite article, the conic regions, i.e., corresponding to the leading and trailing edges of the rotor blade, will be most vulnerable to unacceptable variations in fiber volume. Another disadvantage to filament winding relates to the difficulty associated with establishing fiber orientations at or near 0.degree. relative to the longitudinal or winding axis of the mandrel assembly. Insofar as filament winding apparatus are deficient in this regard, it is common practice to periodically interrupt the winding operation to interleave unidirectional, i.e., 0.degree., fibers.
Reavely et al., U.S. Pat. No. 4,621,980, describes an improved structural rotor blade spar manufactured via a filament winding process. Reavely teaches a multi-layered filament wound composite spar incorporating graphite fiber layers having a fiber orientation below about .+-.35.degree. relative to the longitudinal axis and polyaramide fibers oriented below about .+-.15.degree.. Such orientations produce an acceptable combination of axial and torsional stiffness while enabling the use of a filament winding process. That is, low orientation angles are used to achieve the desired axial stiffness inasmuch as filament winding is not amenable to laying fibers at or near 0.degree. relative to the longitudinal axis. The process, therefore, provides a lay-up which compromises the desired fiber orientation, i.e., at or near 0.degree., to enable the implementation of a filament winding or automated manufacturing process.
The prepreg lay-up technique employs the use of discrete plies or laminates of pre-impregnated composite fabric, which are hand-stacked and interleaved over an inflatable mandrel assembly. The mandrel assembly is placed in a matched metal mold and cured in an autoclave oven for application of heat and pressure. When molding an elliptically shaped article, the lay-up will typically incorporate "slip-planes" located at the interface between overlapping composite plies. These regions of overlap permit the composite lay-up to expand (via slippage across the interface) when curing the composite article. Advantages to the use of a pre-impregnated composite material include ease of compaction, uniform laminate quality, minimal investment for capital equipment, and the ability to selectively orient and build-up material in particular regions. Disadvantages include high labor costs due to the laborious hand lay-up process, comparatively higher probability of operator error (e.g., an operator may inadvertently omit a ply in a multi-ply laminate), and discontinuity of fibers within the laminate.
Salkind et al. U.S. Pat. No. 3,782,856 teaches a twin beam composite spar having a plurality of high tensile strength fibers in a suitable binding matrix. The fibers include off-axis and unidirectionally oriented fibers which are combined to provide high torsional and axial stiffness. High modulus graphite fibers oriented at .+-.45.degree. relative to the longitudinal axis are employed for maximum torsional stiffness. Low modulus fiberglass fibers oriented at 0.degree., i.e., along the longitudinal axis, are incorporated for maximum axial stiffness. Insofar as Salkind intersperses these laminates, and/or bonds packs of off-axis fibers to packs of unidirectional fibers, microcracking will result which adversely affects laminate strength. Furthermore such interspersion of alternating layers of material (one graphite, one fiberglass etc.) is highly labor intensive and prone to operator error.
U.K. Patent Application 2,040,790 discloses a method for manufacturing a constant width composite spar including the steps of: wrapping layers of pre-cut composite fabric about an inflatable mandrel assembly, and forming overlap joints in the upper and lower sidewall regions of the composite spar, i.e., corresponding to the upper and lower airfoil surfaces of the rotor blade. The overlap joints provide slip plane surfaces to facilitate compaction of the composite material during curing operations. Disadvantages to forming overlap joints in the upper and lower sidewalls relates to trapping or pinching the composite fabric upon inflation of the mandrel. Should the inflatable mandrel trap the composite fabric, the slip planes are unable to perform their desired function thereby resulting in poor laminate compaction. This location is also undesirable since the upper and lower sidewall regions are exposed to high flapwise bending stress which results in the superimposition of stresses induced by the joint interface (i.e., stress concentrations).
Kamov, U.S. Pat. No. 3,967,996, and U.K. Patent Applications GB 2040790A and GB 2148 821 are other examples of pre-preg lay-up processes for manufacturing composite spars.
A need, therefore, exists to provide a composite spar structure which optimally blends the desired structural characteristics with a fabrication process which facilitates low cost manufacture. More specifically, a need exists to produce a structurally efficient composite spar which obviates the effects of microcracking, minimizes thermal strain, minimizes weight, and provides damage tolerance. Furthermore, there exists a need to provide a manufacturing process which permits uniform composite material compaction, ease of material handling, reduced hands-on labor during assembly, and produces a composite spar having improved mass distribution properties.